turbine engine component for adaptive cooling

ABSTRACT

A turbine engine component, particularly an aerofoil, including a body is provided. The turbine engine component includes a first surface exposed to a working fluid of high temperatures during operation, a second surface including a depression, the depression is exposed to a cooling fluid during operation and oriented such that, starting from the second surface, the depression deepens in the direction of a back face of the first surface. Furthermore the body includes a body portion between the back face and the first surface. The depression is defined such that a diameter of the depression decreases from the second surface in direction of the back face.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority of European Patent Office applicationNo. 1015248.8 EP filed Feb. 2, 2010, which is incorporated by referenceherein in its entirety.

FIELD OF THE INVENTION

The invention relates to adaptive cooling of a hot turbine enginecomponent, particularly a turbine aerofoil within a turbine section of agas turbine engine.

BACKGROUND OF THE INVENTION

Components such as gas turbine blades, vanes and others in a turbine ora combustion section of a gas turbine may be affected by hot fluidsresulting in oxidation of the components. This problem is addressed inseveral ways, e.g. by using heat resistant materials or by applying acoating, like a thermal barrier coating (TBC). Alternatively oradditionally the components may be cooled. For this additional cooling,cooling air may be supplied and cooling features and cooling holes maybe designed in regions to be cooled.

Even though cooling may be necessary, the existence of cooling holes andthe supply of cooling air to the region to be cooled—the cooling air maytypically be taken from a compressor section of the gas turbine—may havenegative side effects on the overall performance of the compressorand/or of the turbine section.

Thus, a gas turbine may be designed such that only that amount ofcooling air is provided and distributed that is considered necessary. Ininstalled gas turbines, even if thoroughly designed, material may stilloxidise during operation, e.g. due to high temperature peaks.

According to patent application US 2007/0036942 A1 an apparatus foradaptive cooling is disclosed comprising a first component having atleast one aperture extending therethrough with a sacrificial componentpositioned within the at least one aperture. The first component isoperable at a maximum duty temperature and the sacrificial component hasa melting or sublimation point below the maximum duty temperature of thefirst component. The sacrificial component defines an effectiveaperture, the size of which may be increased if, in use, the sacrificialcomponent is subjected to a temperature between the melting orsublimation point of the sacrificial component and the maximum dutytemperature of the first component. The sacrificial component is aseparate component than the first component.

A similar solution is shown in patent U.S. Pat. No. 4,136,516, in whichplugs are provided to initially close secondary air outlets. The plugsare fabricated of a material having a lower melting point temperaturethan that of the remainder of the blade airfoil portion such that, inthe event of failure in the primary cooling system, the plugs will melt,thereby permitting the secondary coolant to rush through the blade andprovide internal cooling thereof. A fairly similar technology isdisclosed in patent application EP 1416225 A1 and patent U.S. Pat. No.3,990,837.

U.S. Pat. No. 7,241,107 B2 shows that cooling air passages may becoated. The coating is being made of a material that has an oxidizingproperty such that the material oxidizes away and opens the passage tomore flow when exposed to a temperature above a critical temperature ofan airfoil. When the airfoil surface is not properly cooled by a flowpassing through the passage, the material oxidizes away until the sizeof the passage increases to allow for the proper amount of cooling airto flow to cool the airfoil. The passages are through holes and eachhave a fixed diameter. To coat these passages may be difficult.

According to patent application GB 2259118 A a leading edge of anairfoil, particularly a coating and airfoil material, may erode orcorrode such that a passage for cooling air—the passage being builtinside of the airfoil—may become exposed to allow a cooling fluid toeffect film cooling of a selected portion of the airfoil.

SUMMARY OF THE INVENTION

The present invention seeks to find a solution in which the time ofoperation can be stretched further and downtime may be avoided even whenoxidation of parts has already occurred, so that repair or replacementof parts can be performed at the next maintenance schedule and stillproviding a mode of operation with only marginal performance losses.

This objective is achieved by the independent claims. The dependentclaims describe advantageous developments and modifications of theinvention.

In accordance with the invention there is provided a turbine enginecomponent, particularly an aerofoil, having a body comprising a firstsurface exposed to a working fluid of high temperatures duringoperation, so that erosion and/or corrosion, e.g. particularlyoxidation, may take place on the first surface. The body is furthercomprising a second surface comprising at least one depression, thedepression being exposed to a cooling fluid during operation, thedepression being oriented such that, starting from the second surface,it deepens in direction of a back face of the first surface. Adepression may be in form of a blind hole or a trench. The body alsocomprises a body portion between the back face and the first surface,wherein a diameter of the depression decreases from the second surfacein direction of the back face. Thus, the depression can be seen as tohave a particularly V-shaped cross section.

The turbine engine component may be, besides the already mentionedaerofoils like blades or vanes in a turbine section of a gas turbineengine, a component inside a hot fluid path or a component defining andguiding the hot fluid path. This component may be located in the turbinesection, in a combustor section of a gas turbine engine or anyintermediate component—like a transition duct between combustor andturbine section—or downstream component—like an exhaust or a diffuser ofthe gas turbine engine that is downstream of the turbine section.

More precisely, the turbine engine component may be any component withina gas turbine engine or a similar type of engine having surfaces overwhich hot gases are swept. Furthermore these surfaces may be part of avane aerofoil and/or a vane platform within a turbine section, and/or ablade aerofoil and/or a blade platform and/or a blade shroud within theturbine section. The surface may also be part of a combustor.Additionally the surface may be part of a static shroud and/or a heatshield. Besides, the invention is also applicable to other hot gas sweptsurfaces downstream of the turbine section in a diffuser and/or in anexhaust, including heat recovery system.

The depression may be casted or machined. If instead of a blind hole athrough hole is produced during manufacturing, the through hole may befilled again partly, covered and/or plugged so that the first surfacemay be closed again.

The through hole may be filled by a heat resistant material, possibly bythe same material as the body, or of a material with a melting pointthat will permanently exceed the temperature during operation. Thus,this heat resistant material is not designed to melt during operation.But the surface directed to the hot fluid path of the heat resistantmaterial may encounter oxidation.

In contrast to prior art disclosure, no special coating material isneeded that is merely applied to be oxidised away faster than theremainder of the first surface. Additionally, the invention is notdirected to oxidise surfaces of cooling passages. It is directed toscenarios in which a hot surface that is designed to guide the hotfluid—the first surface according to the invention—may oxidise.

The invention is advantageous, if the first surface oxidises until thebody portion between the back face and the first surface gets oxidisedaway, so that the depression becomes a passage to the hot fluid path.Due to the fact that the depression is being exposed to the coolingfluid, this newly built passage performs the function of a cooling holevia which the cooling fluid will be guided in direction of the firstsurface. This results in cooling of the first surface and itssurroundings in a proximity of the passage. As a consequence theoxidation may stop in that area.

According to the invention the depression has a specific shape, so thatthe diameter of the depression decreases from the second surface indirection of the back face of the first surface. Once a passage has beenbuilt, this passage has an exit with a first diameter, as defined by thewidth of the depression, providing a first amount of cooling fluid. Thiswill stop or slow down further oxidation in the future. Assuming,oxidation will not be prevented completely because the first amount ofcooling fluid may not be sufficient to do so, the first surfaceincluding the rim of the newly built passage may oxidise further. Thisresults in a further gradual reduction of surface material and also inenlarging the diameter of the cooling passage, due to the specific shapeof the depression with its tilted walls. With the new and largerdimensions of the exit of the passage—having a second diameter—, asecond amount of cooling fluid will be provided to the surroundings ofthe exit. The second diameter will be greater than the first diameter,the second amount of cooling fluid will exceed the first amount ofcooling fluid. This principle will continue until enough cooling fluidwill pass the passage so that no oxidation will occur anymore in thesurroundings of the passage.

To summarise, the inner walls of the passage itself will not oxidise butthe first surface will continue to oxidise which then directly has aneffect on the diameter of the exit of the passage.

Thus, the invention advantageously provides a self adapting principle toadapt to the conditions within the hot fluid path.

In contrast to provide a large number of cooling holes duringmanufacturing to accommodate the needs of worst case scenarios inregards of hot spots, the invention is advantageous because, if the heatwill never reach an oxidising temperature, the depression will neverbecome a cooling passage and therefore no performance losses will occur.

As a consequence, uniform components may be built and installed, eventhough they will be impacted differently by heat due to their differentlocations of installation. Some of the components will operate withoutoxidisation and others may be affected by oxidisation.

In an advantageous embodiment of the invention a part of the bodyadjacent to the body portion is of a further material that is lesssusceptible to erosion and/or corrosion, particularly oxidation, thanthe body portion. The compositions of the material may be seamlesslychange so that this effect may take place. Alternatively the part of thebody adjacent to the body portion may be coated.

The last paragraph is based on the fact that one might anticipate atwhich positions hot spots will occur. This may not always be the case.The invention is also advantageous if the positions of hot spots willnot be known beforehand or on which components the hot spots will occur.This may be the case if the locations of the hot spots depend on how theparts are assembled. In this case a plurality of depressions may bearranged in many places on several components and only these depressionswill become holes eventually that are actually located at or near a hotspot.

The body portion may also incorporate a protective coating over thefirst surface. Cooling holes would then be exposed after the coating andpart of the component material have been lost.

It has to be noted that embodiments of the invention have been describedwith reference to different subject matters. In particular, someembodiments have been described with reference to apparatus type claimswhereas other embodiments have been described with reference to methodtype claims. However, a person skilled in the art will gather from theabove and the following description that, unless other notified, inaddition to any combination of features belonging to one type of subjectmatter also any combination between features relating to differentsubject matters, in particular between features of the apparatus typeclaims and features of the method type claims is considered as to bedisclosed with this application.

The aspects defined above and further aspects of the present inventionare apparent from the examples of embodiment to be described hereinafterand are explained with reference to the examples of embodiment.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments of the invention will now be described, by way of exampleonly, with reference to the accompanying drawings, of which:

FIG. 1: shows schematically an aerofoil with different depressionsaccording to an embodiment of the invention;

FIG. 2A: illustrates a cross section along a plane indicated by a lineA-A in FIG. 1 showing two depressions;

FIG. 2B, 2C, 2D: illustrate in a cross sectional view furtheralternative forms of depressions.

The illustration in the drawing is schematical. It is noted that forsimilar or identical elements in different figures, the same referencesigns will be used.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to FIG. 1, one half of an aerofoil 1 of a turbine is shownin a perspective view as seen from the inside of the aerofoil 1. Thesecond half of the aerofoil 1 that would disallow the view to the insideof the aerofoil 1 is not shown is only indicated by broken and dashedlines.

The aerofoil 1, for example a guide vane or a blade within a turbinesection of a gas turbine engine, represents the turbine engine componentas defined in the invention and will be present in a working fluid flow30 during operation. This working fluid flow 30 is indicated as a doublearrow and may be a hot gas provided by a combustion chamber locatedupstream. The aerofoil 1 has a circumferential wall 2 as a body ofturbine engine component and a cavity to allow cooling fluid—a coolingfluid flow 31 is indicated via a further arrow—to cool the wall 2 fromthe inside and to supply cooling holes (not shown in the figure).

The wall 2 has a first surface 10 which is exposed, in use, to theworking fluid flow 30—a hot gas path—of high temperatures. The firstsurface 10 is directed to the outside of the aerofoil 1.

The wall 2 may be coated from the outside (not shown in FIG. 1;indicated in FIG. 2C), for example by a thermal barrier coating oroxidation resistant coating, such as MCrAIY. In this case the firstsurface 10 may also comprise the coating.

The wall 2 has a second surface 11 which is facing away from the hot gaspath and is being exposed to the cooling fluid flow 31. For the aerofoil1, the second surface 11 is the inner wall of the aerofoil 1,surrounding the cavity.

The second surface 11 comprises a plurality of depressions 20, 21 indifferent forms, e.g. a first depression 20 in form of a truncated cone,having a circular rim with the second surface 11. As a second depression21 a trench is indicated having two side planes and a third bottomplane. Further types of depressions will be shown in FIG. 2.

The depressions 20, 21 are being exposed to the cooling fluid duringoperation. At the time of installation, the depressions 20, 21 are blindholes and may not supply any cooling fluid to the first surface 10.

As already indicated, each of the depressions 20, 21 is oriented suchthat, starting from the second surface 11, the depression 20, 21 deepensin direction of a back face 12 of the first surface 10, and at the sametime, a diameter of the depression 20, 21 decreases from the secondsurface in direction of the back face, thus resulting in a depression inform of a truncated cone (like the first depression 20), a truncatedpyramid, a truncated triangular prism (like the second depression 21),or any kind of frustum, for example.

The back face 12 will be the result of the truncation. The back face 12may be substantially flat.

A diameter may be defined as the distance between two opposing points ona wall of the depression 20, 21. The diameter of the depression 20, 21decreases from the second surface 11 in direction of the back face 12,with a first diameter d1 at the second surface 11 and a second diameterd2 at the back face 12, with d1 greater than d2.

As it should become clear by also consulting the figures, walls of thedepression 20, 21 are tilted, narrowing the diameter until the back face12 is reached, which defines the narrowest plane with the depression 20,21.

Furthermore a body portion 40 can be defined between the back face 12 ofthe first surface 10 and the first surface 10, which is part of the wall2. The body portion 40 may comprise less material than other areas ofthe wall 2 and may define a narrow section of the wall 2. This isintentional, so that in case of oxidation of the first surface 11 of thewall 2, the wall 2 may be breached at this location and will result in apassage (see FIG. 2A and FIG. 2C, reference sign 50) for cooling air.

In FIG. 1 the aerofoil 1 is shown as manufactured. All depression 20, 21may be blind holes which have not become passages yet.

In FIG. 2A the same is shown for a short section of the wall 2 as across section from a direction indicated by the line A-A in FIG. 1.Furthermore FIG. 2A also shows the appearance of the wall section onceoxidation of the first surface 10 has taken place.

According to FIG. 2A, a section of the wall 2 is shown, with its firstsurface 10 that is directed to the hot gas path. Hot streaming gas isindicated again via an arrow for the working fluid flow 30. In thesecond surface 11, as before, the first depression 20 in form of atruncated cone and a second depression 21 in form of a truncatedtriangular prism is shown, which both appear in a cross section astruncated trapezoid or trapezium.

The first depression 20 has an opening of the first diameter d1 thennarrows until the back face 12 is reached with a diameter of the seconddiameter d2. The width of the material of the body portion 40 isindicated by the width w and may be a fraction of a width of a mediumdistance between the first surface 10 and the second surface 11.

The first drawing of FIG. 2A shows the aerofoil 1 without beingoxidised, for example when manufactured or if installed in an area inwhich no extreme temperatures occur.

The second drawing of FIG. 2A show the aerofoil 1 after oxidisation hasoccurred for a longer time span. The first surface 10 is affected byoxidisation resulting in material loss and in an oxidised first surface13. If the temperature remains extremely hot, the oxidisation of theoxidised first surface 11 continues, until all the material of the bodyportion 40 has vanished. In this case the blind holes of the depressions20, 21 will become through holes, i.e. passages 50 which allow coolingair to pass, as indicated via arrows for cooling fluid flow 31. As afurther consequence the cooling air will cool the oxidised first surface13 in a proximity of the passage 50.

This will diminish the oxidisation of the already oxidised first surface13 and slowing down the material degradation at the first surface 10.

If still temperature peaks occur, the material degradation of the firstsurface 10 may continue at a slower pace. This result in furtheroxidisation, as indicated by a dashed line for a further oxidised firstsurface 14. As it can be seen, the height of the passages 50 will alsodecrease, which has a direct impact on the width of the passage 50, dueto the tilted walls of the passage 50. The throughput of cooling airwill increase through the passage 50, because the second diameter d2indicating the diameter of the passage exit will also increase. This hasthe consequence that the oxidisation of the first surface 10 willdecrease further until a stable point of operation is reached, at whichenough cooling air is provided via the passage 50.

Thus, the invention provides a self adjusting principle, in whichcooling passages will be created “automatically” for hot spots and theturbine engine component will adapt the diameters of the exits of thecooling passages “itself” such that the minimum needed cooling air isprovided, so that no oxidisation takes place any further.

Even though this principle is explained for an aerofoil, the inventioncan be applied to other turbine engine components being affected by hotfluids. Dependent on the location of the turbine engine component,different types of depressions with different orientations may beadvantageous. As examples, several types of depressions are shown inFIG. 2B and 2C.

In FIG. 2B a number of depressions 22, 23, 24, 25 are shown, that areslight modifications of the first depression 20 and the seconddepression 21.

The drawing of FIG. 2A showed depressions 20, 21 that show a planesymmetry in regards of a plane through the centre of the back face 12,the plane being perpendicular to the first surface 10 extendingperpendicular to the plane of projection. Due to the symmetry tiltedwalls 51 of the depression 20, 21 have the same length in the projectionof FIG. 2A. Depressions with different walls 51 are shown for thedepressions 22 and 24 in FIG. 2B. This allows to direct the depressionin a wanted direction, so that the cooling fluid, once the passage 50has been occurred, will be guided in a the wanted direction, for exampleto provide film cooling or alternatively to create turbulences.

Depression 23 shows a specific form, in which the depression is a coneor a trapezoid without truncation. Thus, the back face 12 may only be apoint or a line. This will result in a very small passage 50 at thebeginning.

Even though up to now only flat walls of the depressions 20, 21, 22, 23,24 were shown, depression 25 shows a different form, in which the walls51 are convex.

Other forms may be also advantageous.

Not discussed so far, the wall 2 of the aerofoil 1 may be designed suchthat the oxidation mainly takes place or will be affecting the materialfaster on the first surface in the area of the body portion 40. This maybe possible if a part of the wall 2 adjacent to the body portion 40 isof a further material that is less susceptible to erosion and/orcorrosion, e.g. oxidation, than the body portion 40. This allowscreating a passage 50 for cooling air even before the remainder of thefirst surface 10 of the wall 2 is totally oxidised.

This may also be possible, if only parts of the first surface 10 of theaerofoil 1 will be coated by an oxidation resistant coating, leavingcoating free gaps in the area of the body portion 40.

FIG. 2C shows an embodiment in which the complete first surface 10 iscoated. The body portion 40 is sealing the passage 50 to become adepression in form of a blind hole merely by a coating 60. The principleof operation does not differ from the previously said, thus resulting ina passage 50 as shown in the second drawing of FIG. 2C. The solution ofFIG. 2C may be advantageous in respect of manufacturing the aerofoilwith its depressions, because the depression can be manufacturedinitially as through hole and then be closed by the coating 60.

Besides this approach the aerofoil with depressions may be manufacturedby casting the aerofoil including its depressions or by casting a solidaerofoil without depressions and later drilling the depressions.

FIG. 2D shows a slightly different embodiment, in which a depression 27is a blind hole as in FIG. 2A, with an additional coating 60 as knownfrom FIG. 2C. Thus, the cooling passage 50 will then be exposed afterthe coating 60 and part of the wall 2 material have been lost. The bodyportion 40, which may oxidise away, comprises a portion of the coating60 and a portion of the wall 2 between the back face 12 and the firstsurface 10, as indicated in FIG. 2D. As before, in FIG. 2D the part ofthe aerofoil 1 is shown once at the beginning of operation and in asecond drawing after oxidisation has taken place and the passage 50 hasbeen formed.

This invention is particularly advantageous, because the operation timeof turbine engine components that may be affected by oxidisation may beextended. Furthermore the invention is directed to protect installedcomponents in a way that a slight oxidation of surfaces will beaccepted. Besides, an adaption to the needed amount of cooling fluid ispossible, and due to the fact that the walls of the depression aretilted, this adaption may take less time than in other implementations.

Not mentioned so far, the invention may have further advantages inrespect of internal convection cooling. One of the benefits of a tapereddepression is that it increases the internal surface area of the wall towhich it is applied. This means that the depression provides somecooling benefit even before the outside part—the body portion 40—of thedepression is removed and a passage is built. If there is activeimpingement cooling on the inside surface, as there often is in anaerofoil, this cooling benefit could be significant. In this case, todistinguish from prior art convective cooling solutions, it may be thetarget to design the depression such that in a perfect installation theconvective cooling may be sufficient and no erosion and/or corrosiontakes place. In a real life installation, in which hot spots may arisedue to assembly and/or manufacturing inaccuracies, the body portion 40may be sized such that there is a chance that a passage 50 may ariseduring typical time of operation of the component. Thus, the bodyportion 40 may have a lesser width than known depressions that aresolely intended for convection cooling.

Besides, the convective heat transfer coefficient will also be enhanceddue to turbulences of the cooling air resulting of the tapereddepression.

In other words, the depression may be arranged such that the bodyportion may be eroded and/or corroded completely and generating thepassage 50 before a next scheduled inspection of the gas turbine incases of continuously occurring hot spots.

Besides, it has to be mentioned that the invention allows easymanufacturing. Particularly due to the tapered form of the depressions,the turbine engine component may be easier to cast. A ceramic formingthe depression during the casting process is stronger with a taperedshape than with a parallel sided shape.

Even though the invention was explained for depressions with acontinuous reduction in the diameter of the depression, alsonon-continuous forms are possible. For example a depression could startfrom the second surface with a cylindrical shape, later followed by atapered portion in direction of the back face.

1.-10. (canceled)
 11. A turbine engine component including a body,comprising: a first surface exposed to a working fluid having a hightemperature during operation; a second surface including a depression,the depression being exposed to a cooling fluid during operation andoriented such that, starting from the second surface, the depressiondeepens in a direction of a back face of the first surface; and a bodyportion disposed between the back face and the first surface, wherein adiameter of the depression decreases from the second surface in thedirection of the back face.
 12. The turbine engine component accordingto claim 11, wherein the body portion is comprised of a material and/orof a width between the first surface and the back face of the firstsurface, and wherein when the body portion is exposed to a specifictemperature, erosion and/or corrosion on the first surface side of thebody portion a passage through the depression will open.
 13. The turbineengine component according to claim 12, wherein a part of the bodyadjacent to the body portion includes a further material that is lesssusceptible to erosion and/or corrosion than the body portion.
 14. Theturbine engine component according to claim 13, wherein the furthermaterial is an oxidation resistant coating.
 15. The turbine enginecomponent according to claim 12, wherein the open passage provides acooling fluid producing a cooling to a proximity of the first surface.16. The turbine engine component according to claim 15, wherein thecooling is film cooling.
 17. The turbine engine component according toclaim 15, wherein the diameter of the depression decreases from thesecond surface in the direction of the back face such that, when in use,continuing erosion and/or corrosion on the first surface side of thebody portion results in widening the passage to let pass a larger amountof cooling fluid.
 18. The turbine engine component according to claim17, wherein the diameter of the depression decreases from the secondsurface in direction of the back face such that, when in use, continuingwidening of the passage due to erosion and/or corrosion provides anincreasing amount of passed cooling fluid such that a temperature of aproximity of the first surface is reduced to a value at which erosionand/or corrosion on the first surface side of the body portionterminates.
 19. The turbine engine component according to claim 11,wherein the turbine engine component is an aerofoil.
 20. A gas turbineengine, comprising: a turbine engine component, comprising: a firstsurface exposed to a working fluid having a high temperature duringoperation, a second surface including a depression, the depression beingexposed to a cooling fluid during operation and oriented such that,starting from the second surface, the depression deepens in a directionof a back face of the first surface, and a body portion disposed betweenthe back face and the first surface, wherein a diameter of thedepression decreases from the second surface in the direction of theback face.
 21. The gas turbine engine according to claim 20, wherein theturbine engine component is disposed in a vane aerofoil, and/or a vaneplatform, and/or a blade aerofoil, and/or a blade platform, and/or ablade shroud, and/or a transition duct between a combustor and a turbinesection, and/or a combustor, and/or a static shroud, and/or a heatshield, and/or a diffuser, and/or an exhaust.
 22. The gas turbine engineaccording to claim 20, wherein the body portion is comprised of amaterial and/or of a width between the first surface and the back faceof the first surface, and wherein when the body portion is exposed to aspecific temperature, erosion and/or corrosion on the first surface sideof the body portion a passage through the depression will open.
 23. Thegas turbine engine according to claim 22, wherein a part of the bodyadjacent to the body portion includes a further material that is lesssusceptible to erosion and/or corrosion than the body portion.
 24. Thegas turbine engine according to claim 23, wherein the further materialis an oxidation resistant coating.
 25. The gas turbine engine accordingto claim 22, wherein the open passage provides a cooling fluid producinga cooling to a proximity of the first surface.
 26. The gas turbineengine according to claim 25, wherein the cooling is film cooling. 27.The gas turbine engine according to claim 25, wherein the diameter ofthe depression decreases from the second surface in the direction of theback face such that, when in use, continuing erosion and/or corrosion onthe first surface side of the body portion results in widening thepassage to let pass a larger amount of cooling fluid.
 28. A method foradaptive cooling a turbine engine component, the method comprising:providing a cooling fluid to the turbine engine component including abody, the gas turbine component comprising: a first surface exposed to aworking fluid of a high temperature during operation, a second surfacecomprising a depression, the depression being exposed to the coolingfluid during operation and oriented such that, starting from the secondsurface, the depression deepens in a direction of a back face of thefirst surface, and a body portion disposed between the back face and thefirst surface; and applying heat to the first surface such that amaterial of the first surface is eroded and/or corroded on the firstsurface side of the body portion so that a passage to the depressionwill open allowing the cooling fluid to pass, wherein a diameter of thedepression decreases from the second surface in direction of the backface.
 29. The method according to claim 20, wherein the depression iscast or machined.